The present invention refers in general to a leading edge for aircraft lifting or supporting surfaces, such as wings and stabilizers. One object of the present invention is to provide an optimized structure for a leading edge of an aircraft, in order to reduce its weight and to thereby reduce fuel consumption.
Additionally, it is also an object of the present invention to provide a leading edge of an aircraft which can be manufactured with a reduced number of components, in order to simplify its assembly and manufacturing process.
Aircraft lifting surfaces such as wings, Horizontal Tail Planes (HTP), Vertical Tail Planes (VTP), etc., are formed by skin panels reinforced internally by a supporting structure, which typically comprises longitudinal front and rear spars, transverse ribs joining the spars, and stringers between the ribs and skin panels.
On the leading edge of a wing, that is, on the front edge of the wing as seen in the direction of flight, there are one or more leading edge sections longitudinally arranged to form the outermost surface of the wing. The leading edge is coupled with the torsion box of the wing and comprises its own skin panel and support structure.
Known leading edge designs comprise skin panels internally stiffened by several leading edge ribs. In the case of large commercial aircraft, two additional metallic spars are used, wherein one of them is vertically arranged next to the foremost point or nose of the leading edge, and the other one is diagonally arranged in a cross-sectional view of the leading edge.
Conventionally, the skin panels of leading edges have a C-shaped configuration in a cross-sectional view, and are constructed as non-monolithic structures which consist of a sandwiched structure formed by a honeycomb core with Carbon Fiber Reinforced Plastic (CFRP) or Glass Fiber Reinforced Plastic (GFRP) facesheets, or by a honeycomb core covered by a metallic protection sheet made of steel or aluminium alloys to protect the leading edge from erosion phenomena. U.S. Pat. No. 6,616,101B2 is an example of a leading edge having sandwich skin panels of the above-mentioned type.
It is known that honeycomb sandwiched panels are difficult to repair and suffer from water ingestion during flight.
Additional elements such as leading edge extensions are used in transition areas between the leading edge and the fuselage. These elements are usually constructed with the same sandwich philosophy used for skin panels.
Due to the large span of most leading edges, these are split in several sections which are constructed as sub-assemblies to ease their manufacture, assembly and maintenance. Although the profile of the leading edge varies progressively from root to tip, the internal supporting structure is basically the same in all the sections of the leading edge.
There is therefore the need for leading edge structures which are lighter and which can be constructed with a reduced number of components in order to simplify their manufacture.